The low earth orbit of a spacecraft is surrounded by the "plasma" of which space is composed, space not being quite empty, and being even less empty close to objects such as earth having a gravitational force. The plasma, also called space plasma, is composed of neutrally charged particles having no electrical charge, negatively charged electrons, and various positively charged ionic species. The positive ionic species are much more massive on an atomic scale than electrons, and therefore possess a lower thermal velocity. Even though the plasma is neutral, the electron current density in a particular direction is greater than the ion current density. As a result, when a spacecraft is immersed in the low earth orbit plasma, the exterior surface of the spacecraft will tend to accumulate a negative charge. As the spacecraft accumulates charges, the resulting negative electric field begins to repel electrons while attracting the positive ions. Charging continues until an equilibrium potential is reached, resulting in no net current being collected. In low earth orbit, with no voltage potential created inside the spacecraft to change the exterior surface voltage potential, the equilibrium potential is roughly a few volts negative with respect to the surrounding space plasma potential.
Negative charge, even if only a few volts negative, attracts the relatively massive positively charged ionic species. At low voltages, that is not too much of a problem. However, advanced spacecraft have higher power demand. Power and energy are a function of volts times amperage. Wattage=volts x amps. In order to increase power without increasing the voltage, one solution is to increase the amperage.
The difficulty with increasing amperage, as explained below, is that it adds weight and mass and generates heat. If the amperage is increased to meet increased internal power demand of the spacecraft, the size of wire and electrical connection must be increased, adding to the weight of a spacecraft. Each pound of increased weight of a spacecraft increases launch cost.
Because of launch cost, there is a premium on volume and hundreds of electrical and electronic components are placed into a spacecraft because of the launch cost. For higher wattage (which is equal to amps squared times resistance) without higher voltage, there must be higher amperage, which generates unacceptable heat from the electrical loads in the spacecraft. The higher amperage tends to generate more heat in the electrical components, and means that overheating of close packed electrical components is more likely to occur, requiring more spacing of components in contradiction to the objective of maximizing the use of volume.
Further, the current carrying capacity of a wire of given cross-sectional area is limited because a smaller wire has a higher resistance per given linear measure. Thus to increase the amperage, the cross-section of the wire must be increased which is undesirable because of the added mass of the wiring and wiring harness.
The alternative solution to meet the higher power demand is to increase the voltage. If the system could be operated at higher voltage for a given power demand, such higher voltage would reduce the amperage requirement and the weight of the spacecraft.
Power for a spacecraft, on a long-term basis, is generated using multiple solar cells in solar arrays. Each solar cell generates a small voltage potential between two terminals, a positive and a negative terminal, and generates a small direct current measured in amps. To achieve high voltages, the solar cells are linked in series, as in a string, with the terminals at each end, "adding" the voltages. To increase the amperage without changing the voltage, the strings of solar cells are linked in parallel which increases the available amperage at the terminals of the solar cell array.
Higher array voltages, unfortunately, intensify the electromagnetic interaction between the spacecraft and the space plasma under present spacecraft designs. A spacecraft in low earth orbit can experience many adverse effects on the solar arrays and the spacecraft. These effects include sputtering of spacecraft surfaces, contamination of the surfaces and arcing. The magnitude of these effects and areas affected on the exterior of the spacecraft depends mainly on the voltage level of the solar arrays and on the grounding configuration of the power system. The sputtering effect is one of erosion of all spacecraft materials exposing underlying regions to atomic oxygen attack. A power system which is negatively grounded to the hull can also suffer arcing effects between the plasma and underlying negatively charged spacecraft surfaces if the dielectric strength of the insulating material for the power feed from a solar array is exceeded. If arcing occurs, pinholes in the insulating material surface occurs exposing the structure to the space plasma, which causes enhanced sputtering to the hull because plasma ions are focused to the pinhole edges by the large electric fields. Material that escapes from the sputter surface is a source of contamination to the spacecraft and in particular affects the solar arrays and certain thermal surfaces. This phenomenon is described and published in NASA Technical Memorandum 103717, "Findings of the Joint Workshop on Evaluation of Impacts of Space Station Freedom Ground Configurations," D. C. Ferguson, D. B. Snyder, and R. Carruth (Nat'l Tech. Inf. Serv., Springfield, Va. 1990). The significance of the sputtering problem was also discussed in presentation in NASA "Sputtering and SSF Implications, D. C. Ferguson, Aerospace Technology Research Project Review, Mar. 6, 1991.
One possible solution is to positively ground the spacecraft hull or containing means from the positive terminal of the solar arrays, and then to link the positive side of the internal electrical components of the spacecraft to the positive hull. This is done, to put it simplistically, by simply connecting the positively charged terminal of the solar array(s) to the spacecraft hull or containing means.
Positively grounding the hull to the positive terminal does create differential potentials between the exposed spacecraft surfaces and the space plasma. They are too small, however, to cause concern. The exposed surfaces of the spacecraft will charge to a slightly positive potential compared to plasma potential. These small positive potentials repel massive positive ions that might cause sputtering.
The difficulty with this approach is that a positive ground architecture is not conventional for modern electronic technology. The choices among semiconductor p-channel devices normally associated with positive ground circuits are rather limited. The range of power, current, and voltage handling capabilities in available p-channel devices (devices having a positively charged case) are not nearly so broad as the range in n-channel devices which have broad availability (having a negatively charged case normally associated with negatively grounded circuits). Circuits using n-channel devices referred to positive ground tend to be rather complex resulting in attendant reliability and mass and weight penalties. Thus, a negative ground electrical system and spacecraft structure is preferable from the perspective of the designer of the internal electrical loads.
One solution to enable the use of a positively grounded hull with negatively grounded electronic devices which is known in the art is to use isolated transformer power architectures. A transformer permits the reversing of the voltage potential with relatively little loss of power for an isolated device. However, a transformer has a certain bulk and mass associated with it. A transformer uses a coupled inductance of at least two induction coils wound around and arranged apart on a metal, usually iron-based core, the "input" coil being called the primary coil or inductor and the "output" coil or inductor being called the secondary coil. In order for a transformer to work, the direct current from the solar cells must be converted to alternating current, which drives the primary coil, and then the output of the secondary coil must be rectified back into direct current. Transformers are relatively heavier and require complicated packaging of the power system and the electrical loads.
For Space Station Freedom, after the plasma effects of the planned 167V system had been confirmed, two remedial solutions were proposed to positively charge the spacecraft. One solution was to ground Space Station Freedom to the positive end of the solar array, but redesign of a number of important components, including the power conversion unit(s), to yield a negatively grounded system was required.
Ultimately, an alternative remedial solution was proposed in the form of a plasma contactor capable of pulling the highly negative spacecraft hull and structure up to the plasma potential. This was to be accomplished by building up a large effective electron emitting area, in the form of a plasma sheath generated by the plasma contactor. The plasma contactor is an active device expending xenon gas in its operation, and capable of pulling the entire solar array positive with respect to the space plasma.
This solution, aside from its complexity and redesign requirements of the power conversion unit(s), has costs in several ways. The currents collected the plasma, especially during times of high solar activity, will be a significant power drain on the system. Redundancy, and some control are required to ensure continuous operation. Contactor replacement or xenon gas replenishment will be necessary after a few years of operation. However, this is believed less expensive than redesigning the electronic components, the electrical system and overall design of the space station at this advanced stage of development.